Electric propulsion has been considered for powering spacecraft and other vehicles because of its potentially high specific impulse which reduces the amount of propellant and therefore system weight for a mission. A high specific impulse, resulting from a high exhaust velocity, requires high power per unit thrust. Therefore, a trade needs to be made between specific impulse, propellant weight, and weight of the power source, whether it is solar photovoltaic, solar thermal, or nuclear. A critical part of electric thrust generation is the process efficiency that relates to both propellant and its power generation system weight. Specifically, this is the efficiency of converting electrical power into thrust by accelerating the propellant. The process is typically not very efficient due to a number of factors including: losses to ionization and/or dissociation that is not recoverable as thrust; high-temperature operation and related radiative losses; required heating of cathode, electron emitter, or accelerating electrodes in the case of different types of ion engines; and arc losses including resistive, radiative and thermal, amongst others. Because of these issues and particularly the efficiency and power system weight, electric propulsion has limited applications, even with its potential for extremely high specific impulse as compared to chemical systems. The current non-chemical thrusters technology includes electro-thermal with a specific impulse of 800 seconds, arc jets with a specific impulse of 900-1,200 seconds, MPD/MHD and Hall-Effect with a specific impulse range from 1,200-2,500 seconds, and electrostatic/ion thrusters from 1,000-25,000 seconds. Electric thrusters depend on electrical energy fed into the reaction mass resulting in lower exhaust velocity and specific impulse, and increased system power-to-thrust ratio.
A system and method for propulsion is needed to increase exhaust velocity, to increase specific impulse, and to decrease system power-to-thrust ratio.